1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to turbine airfoils with a cooling circuit.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine is a very efficient way for converting combustion into mechanical energy used to produce electrical power. A gas turbine engine includes a compressor to compress air, a combustor to mix the compressed air with a fuel and generate a hot gas flow, and a turbine to receive the hot gas flow and drive the turbine shaft. A typical turbine in an industrial gas turbine engine (IGT) will use four stages of stator vanes and rotor blades to progressively convert the energy of the hot gas flow into mechanical energy. A turbine has a temperature operating limit based upon the hottest temperature that the first stage vanes and blades can withstand without damage. The engine efficiency can be increased by increasing the hot gas flow into the turbine. It is therefore desirable to allow for a higher gas flow temperature in the turbine to produce more power using less fuel.
One method of increasing the efficiency of the engine is to provide for internal air cooling of the first stage vanes and blades. Even though the materials have not changed, the air cooled airfoils (blades and vanes) will allow for a higher temperature flow and therefore an increase in the engine efficiency. The cooling circuit includes internal channels and cavities for conductive cooling of the blade and film cooling holes on the airfoil surface that provide a blanket of cooling air between the hot gas flow and the airfoil surface. in film cooling, the cooling air must be channeled through the airfoils with a high enough pressure to prevent blowback ingestion of the hot gas flow through the film cooling holes, and also avoid excessive pressure drop across the film cooling holes which would tend to separate the film of cooling air from the outer surface of the airfoil which would degrade the film cooling effectiveness.
On a turbine blade with a pressure side and a suction side, certain surface areas require film cooling while others can make due with the convective cooling from the flow of cooling air in the through path channel such as a leg of the serpentine flow circuit. this is especially true for the first and second legs of the serpentine flow circuit, since the cooling air entering these channels is fresh air that have not been heated too much.
Another method of improving the engine efficiency is to use less cooling air in the airfoils to provide the same amount of cooling. The compressed air used as the cooling air is typically air bled off from the compressor. Energy is required to compress the cooling air, and therefore energy is lost and the engine efficiency is lowered. Complex internal air cooling circuitry has been proposed to provide a maximum amount of cooling while using a minimum amount of cooling air. The locations of film cooling holes are strategically placed to provide film cooling to hot spots on the airfoil walls. Cooling air pressures are regulated due to different external flow pressures over the airfoil walls. The external pressure is higher on the pressure side than it is on the suction side, while the hottest region on the airfoil surface appears on the suction side than on the pressure side. Thus, there is always a desire to improve on the prior art internal airfoil cooling air circuitry to provide the maximum amount of cooling while using the minimum amount of cooling air.
U.S. Pat. No. 6,705,836 B2 issued to Bourriaud et al on Mar. 16, 2004 and entitled GAS TURBINE BLADE COOLING CIRCUITS discloses a turbine blade with multiple serpentine flowing cooling circuits separate from one another. One serpentine circuit is on the pressure side of the mid-chord portion, a second serpentine flow circuit is in the trailing edge region, a third serpentine flow circuit is on the suction side at the mid-chord portion, and a central cooling supply channel is between the pressure side and suction side serpentine flow circuits and supplies cooling air to the showerhead arrangement.
U.S. Pat. No. 6,039,537 issued to Scheurlen on Mar. 21, 2000 entitled TURBINE BLADE WHICH CAN BE SUBJECTED TO A HOT GAS FLOW discloses a turbine blade with a series of cooling channels extending from the leading edge region to the trailing edge region, each channel extending from the pressure side wall to the suction side wall to provide near wall cooling for both the pressure and suction sides. One of these channels includes film cooling holes extending onto the pressure side wall and the suction side wall of the blade. One problem with this particular design is that the cooling air supply pressure for the suction side film cooling holes is the same pressure as the pressure side film cooling holes. since the external pressure on the pressure side is higher than the external pressure on the suction side, either too much cooling air is discharged out the suction side film cooling holes or too little discharged out the pressure side film cooling holes. Either way, the cooling of the blade is either too little or uses too much cooling air.
U.S. Pat. No. 5,660,524 issued to Lee et al on Aug. 26, 1997 and entitled AIRFOIL BLADE HAVING A SERPENTINE COOLING CIRCUIT AND IMPINGEMENT COOLING discloses a turbine blade with three separate cooling circuit that include a 2-pass serpentine flow circuit in the leading edge in which the second leg impinges cooling air onto a leading edge cavity connected to a showerhead arrangement of film cooling holes, a trailing edge cooling supply channel that is a single pass channel and connected to exit cooling holes along the trailing edge of the blade, and a 3-pass serpentine flow circuit with a first leg adjacent to the trailing edge cooling supply channel, a second leg forward of the first, and the third leg in the middle of the blade adjacent to the leading edge cooling circuit. The third leg provides impingement cooling to a pressure side impingement cavity and a suction side impingement cavity, with each of the impingement cavities having film cooling holes discharging cooling air onto the blade wall.
U.S. Pat. No. 6,206,638 B1 issued to Glynn et al on Mar. 27, 2001 and entitled LOW COST AIRFOIL COOLING CIRCUIT WITH SIDEWALL IMPINGEMENT COOLING CHAMBERS discloses a turbine blade with a 3-pass (triple pass) serpentine flow cooling circuit extending along the suction side wall and flowing in an aft to-forward direction, and in which each of the legs in the serpentine flow circuit impinges onto an impingement cavity located on the pressure side wall or the leading edge of the blade. Each impingement cavity includes film cooling holes.
U.S. Pat. No. 5,498,133 issued to Lee on Mar. 12, 1996 and entitled PRESSURE REGULATED FILM COOLING discloses a turbine airfoil such as a vane or a blade with two serpentine flow cooling circuit that share a common first leg channel, one flowing in the aft direction and the other flowing in the forward direction, and each channel is connected to an impingement cavity by a metering hole, and the cavities include film cooling holes.
None of the above cited prior art references anticipate nor make obvious the present invention in which a multiple pass serpentine flow cooling circuit provides convective cooling to surfaces of the airfoil on both the pressure side and the suction side that does not require film cooling as well as impingement cooling cavities with film cooling holes on surfaces of the airfoil on both sides that require film cooling while using a minimal amount of cooling air in order to increase the efficiency of the gas turbine engine.
It is therefore an object of the present invention to provide for a cooling air circuit for a turbine airfoil that provides increased cooling while using minimal amount of cooling air in order to increase the efficiency of the gas turbine engine.
It is another object of the present invention to provide for a cooling circuit within a turbine airfoil that can regulate the pressure and amount of cooling air flow in individual areas of the airfoil in order to provide adequate cooling without over-cooling certain areas.